Compound engine assembly with offset turbine shaft, engine shaft and inlet duct

ABSTRACT

A compound engine assembly with an inlet duct, a compressor, an engine core including at least one internal combustion engine, and a turbine section including a turbine shaft configured to compound power with the engine shaft. The turbine section may include a first stage turbine and a second stage turbine. The turbine shaft and the engine shaft are parallel to each other. The turbine shaft, the engine shaft and at least part of the inlet duct are all radially offset from one another. A method of driving a rotatable load of an aircraft is also discussed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.14/864,096, filed Sep. 24, 2015, which claims priority from U.S.provisional application No. 62/118,914 filed Feb. 20, 2015, the entirecontents of which are incorporated by reference herein.

TECHNICAL FIELD

The application relates generally to compound engine assemblies and,more particularly, to supercharged or turbocharged compound engineassemblies used in aircraft.

BACKGROUND OF THE ART

Compound engine assemblies including a compressor used as a superchargeror turbocharger may define a relatively bulky assembly which may bedifficult to fit into existing aircraft nacelles, thus creating somedifficulty in adapting them for aircraft applications.

SUMMARY

In one aspect, there is provided a compound engine assembly comprising:an inlet duct having an inlet in fluid communication with ambient airaround the compound engine assembly; a compressor having an inlet influid communication with the inlet duct, the compressor including atleast one compressor rotor connected to a turbine shaft; an engine coreincluding at least one internal combustion engine in driving engagementwith an engine shaft, the engine core having an inlet in fluidcommunication with an outlet of the compressor; a turbine section havingan inlet in fluid communication with an outlet of the engine core, theturbine section including at least one turbine rotor connected to theturbine shaft, the turbine shaft configured to compound power with theengine shaft; wherein the turbine shaft and the engine shaft areparallel to each other; and wherein the turbine shaft, the engine shaftand a longitudinal central axis of at least part of the inlet duct areall radially offset from one another.

In another aspect, there is provided a compound engine assemblycomprising: an inlet duct having an inlet in fluid communication withambient air around the compound engine assembly; a compressor having aninlet in fluid communication with the inlet duct; an engine coreincluding at least one internal combustion engine in driving engagementwith an engine shaft, the engine core having an inlet in fluidcommunication with an outlet of the compressor; a turbine sectionincluding a first stage turbine having an inlet in fluid communicationwith the outlet of the engine core and a second stage turbine having aninlet in fluid communication with an outlet of the first stage turbine,each of the first stage turbine and the second stage turbine includingat least one rotor connected to a turbine shaft, the turbine shaft andthe engine shaft being in driving engagement with one another; whereinthe turbine shaft and the engine shaft are parallel to each other; andwherein the turbine shaft, the engine shaft at least part of the inletduct are all radially offset from one another.

In a further aspect, there is provided a method of driving a rotatableload of an aircraft, the method comprising: directing ambient air fromoutside of the compound engine assembly into the compound engineassembly through an inlet duct; directing the air from the inlet duct toan inlet of a compressor; directing compressed air from an outlet of acompressor to an inlet of at least one internal combustion engine of acompound engine assembly; driving rotation of an engine shaft with theat least one combustion engine; driving rotation of a turbine shaft of aturbine section of the compound engine assembly by circulating anexhaust of the at least one internal combustion engine to an inlet ofthe turbine section; and compounding power from the turbine shaft withthat of the engine shaft to drive the rotatable load; wherein theturbine shaft and the engine shaft are parallel to each other andradially offset with respect to each other; and wherein the air iscirculated through the inlet duct along a path radially offset from theturbine shaft and from the engine shaft.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic view of a compound engine assembly in accordancewith a particular embodiment;

FIG. 2 is a cross-sectional view of a Wankel engine which can be used ina compound engine assembly such as shown in FIG. 1, in accordance with aparticular embodiment;

FIG. 3 is a schematic tridimensional view of the compound engineassembly of FIG. 1 in accordance with a particular embodiment;

FIG. 4 is a schematic side view of the compound engine assembly of FIG.3, with an engine mount in accordance with a particular embodiment;

FIG. 5 is a schematic cross-sectional side view of the compound engineassembly of FIG. 3, with an inlet duct and firewall according to aparticular embodiment;

FIG. 6 is a schematic front view of the compound engine assembly of FIG.3, according to a particular embodiment;

FIG. 7 is a schematic view of a compound engine assembly in accordancewith another particular embodiment;

FIG. 8 is a schematic tridimensional view of the compound engineassembly of FIG. 7 in accordance with a particular embodiment;

FIG. 9 is a schematic cross-sectional side view of the compound engineassembly of FIG. 8, with an inlet duct and firewall according to aparticular embodiment;

FIG. 10 is a schematic tridimensional view of the compound engineassembly of FIG. 8, with an engine mount in accordance with a particularembodiment;

FIG. 11 is a schematic, exploded end view of a compound engine assemblyin accordance with another particular embodiment; and

FIG. 12 is a schematic side view of part of the compound engine assemblyof FIG. 11.

DETAILED DESCRIPTION

Referring to FIG. 1, a compound engine assembly 10 is generally shown,including a liquid cooled heavy fueled multi-rotor rotary engine core12. The engine core 12 has an engine shaft 16 driven by the engine core12 and driving a rotatable load, which is shown here as a drive shaft 8.The drive shaft 8 may be an integral part of the engine shaft 16, bedirectly connected thereto, or be connected thereto through a gearbox(not shown). It is understood that the compound engine assembly 10 mayalternately be configured to drive any other appropriate type of load,including, but not limited to, one or more generator(s), propeller(s),accessory(ies), rotor mast(s), compressor(s), or any other appropriatetype of load or combination thereof.

The compound engine assembly 10 is configured as a single shaft engine.The term “single shaft” is intended herein to describe a compound enginewhere all the rotating components (compressor rotor(s), turbinerotor(s), engine shaft, accessories) are mechanically linked together,either directly or through one more gearbox(es). Accordingly, a “singleshaft” engine may include two or more mechanically linked shafts. Theterm “single shaft” is intended to be in contrast to an engine havingtwo or more spools which are free to rotate with respect to one anothersuch as to include one or more free turbine(s).

The compound engine assembly 10 includes a compressor 14 feedingcompressed air to the inlet of the engine core 12 (corresponding to orcommunicating with the inlet port of each engine of the engine core 12).The engine core 12 receives the pressurized air from the compressor 14and burns fuel at high pressure to provide energy. Mechanical powerproduced by the engine core 12 drives the engine shaft 16. Each engineof the engine core 12 provides an exhaust flow in the form of exhaustpulses of high pressure hot gas exiting at high peak velocity. Theoutlet of the engine core 12 (corresponding to or communicating with theexhaust port of each engine of the engine core 12) is in fluidcommunication with an inlet of a turbine section 18, and accordingly theexhaust flow from the engine core 12 is supplied to the turbine section18. The turbine section 18 drives the compressor 14 and compounds powerwith the engine shaft 16.

In a particular embodiment, the compound engine assembly includes four(4) major modules: a core module 22 including the engine core 12, agearbox module 20, a cold section or compressor module 24 including thecompressor 14 and a hot section or turbine module 28 including theturbine section 18. In a particular embodiment, the turbine module 28and compressor module 24 are removable by typical maintenance personnel,in the field, with the compound engine assembly 10 remaining attached tothe aircraft, for ease of maintenance, repair and/or replacement. In aparticular embodiment, each of the turbine module 28, compressor module24 and core module 22 can be detached and removed from the compoundengine assembly 10 in an individual and separate manner, i.e. withoutthe need to detach/remove any of the other modules; in a particularembodiment, the components of each module are thus contained in and/ormounted to a casing which defines an enclosure independently of that ofthe other modules. In a particular embodiment, the modularity of thecompound engine assembly 10 may allow reducing or minimizing the numberof parts in the compound engine assembly 10 and/or may enable eachmodule to run at speeds corresponding to optimum performance conditions.

Referring to FIG. 3, the core module 22 includes the engine core 12 anda fuel distribution system 13. In the embodiment show, the engine core12 includes a plurality of rotary engines 12′ drivingly engaged to theshaft 16, and the fuel distribution system 13 includes a common rail 13′feeding a pilot and a main injector for each rotary engine. Although theengine core 12 is depicted as including two rotary engines 12′, it isunderstood that in another embodiment, the engine core 12 may includemore than two rotary engines 12′ (e.g. 3 or 4 rotary engines), or asingle rotary engine 12′. Each rotary engine 12′ has a rotor sealinglyengaged in a respective housing, with each rotary engine 12′ having anear constant volume combustion phase for high cycle efficiency. In theembodiment shown, each rotary engine 12′ is a Wankel engine.

Referring to FIG. 2, an exemplary embodiment of a Wankel engine whichmay be used as rotary engine 12′ in the engine core 12 is shown. EachWankel engine 12′ comprises a housing 32 defining an internal cavitywith a profile defining two lobes, which is preferably an epitrochoid. Arotor 34 is received within the internal cavity. The rotor defines threecircumferentially-spaced apex portions 36, and a generally triangularprofile with outwardly arched sides. The apex portions 36 are in sealingengagement with the inner surface of a peripheral wall 38 of the housing32 to form three working chambers 40 between the rotor 34 and thehousing 32.

The rotor 34 is engaged to an eccentric portion 42 of the shaft 16 toperform orbital revolutions within the internal cavity. The shaft 16performs three rotations for each orbital revolution of the rotor 34.The geometrical axis 44 of the rotor 34 is offset from and parallel tothe axis 46 of the housing 32. During each orbital revolution, eachchamber 40 varies in volume and moves around the internal cavity toundergo the four phases of intake, compression, expansion and exhaust.

An intake port 48 is provided through the peripheral wall 38 forsuccessively admitting compressed air into each working chamber 40. Anexhaust port 50 is also provided through the peripheral wall 38 forsuccessively discharging the exhaust gases from each working chamber 40.Passages 52 for a glow plug, spark plug or other ignition element, aswell the fuel injectors are also provided through the peripheral wall38. Alternately, the intake port 48, the exhaust port 50 and/or thepassages 52 may be provided through an end or side wall 54 of thehousing; and/or, the ignition element and a pilot fuel injector maycommunicate with a pilot subchamber (not shown) defined in the housing32 and communicating with the internal cavity for providing a pilotinjection. The pilot subchamber may be for example defined in an insert(not shown) received in the peripheral wall 38.

In the embodiment of FIG. 3, the fuel injectors are common rail fuelinjectors, and communicate with a source of Heavy fuel (e.g. diesel,kerosene (jet fuel), equivalent biofuel), and deliver the heavy fuelinto the engine(s) such that the combustion chamber is stratified with arich fuel-air mixture near the ignition source and a leaner mixtureelsewhere.

Referring back to FIG. 2, for efficient operation the working chambers40 are sealed, for example by spring-loaded apex seals 56 extending fromthe rotor 34 to engage the peripheral wall 38, and spring-loaded face orgas seals 58 and end or corner seals 60 extending from the rotor 34 toengage the end walls 54. The rotor 34 also includes at least onespring-loaded oil seal ring 62 biased against the end wall 54 around thebearing for the rotor 34 on the shaft eccentric portion 42.

Each Wankel engine provides an exhaust flow in the form of a relativelylong exhaust pulse; for example, in a particular embodiment, each Wankelengine has one explosion per 360° of rotation of the shaft, with theexhaust port remaining open for about 270° of that rotation, thusproviding for a pulse duty cycle of about 75%. By contrast, a piston ofa reciprocating 4-stroke piston engine typically has one explosion per720° of rotation of the shaft with the exhaust port remaining open forabout 180° of that rotation, thus providing a pulse duty cycle of 25%.

In a particular embodiment which may be particularly but not exclusivelysuitable for low altitude, each Wankel engine has a volumetric expansionratio of from 5 to 9, and operates following the Miller cycle, with avolumetric compression ratio lower than the volumetric expansion ratio,for example by having the intake port located closer to the top deadcenter (TDC) than an engine where the volumetric compression andexpansion ratios are equal or similar. Alternately, each Wankel enginemay operate with similar or equal volumetric compression and expansionratios.

It is understood that other configurations are possible for the enginecore 12. The configuration of the engine(s) 12′ of the engine core 12,e.g. placement of ports, number and placement of seals, number of fuelinjectors, etc., may vary from that of the embodiment shown. Inaddition, it is understood that each engine 12′ of the engine core 12may be any other type of internal combustion engine including, but notlimited to, any other type of rotary engine, and any other type ofnon-rotary internal combustion engine such as a reciprocating engine.

Referring back to FIG. 1, in a particular embodiment the compressor 14is a centrifugal compressor with a single impeller 14′. Otherconfigurations are alternately possible. The compressor 14 may besingle-stage device or a multiple-stage device and may include one ormore rotors having a circumferential array of radial, axial or mixedflow blades.

Referring to FIG. 3, the gearbox module 20 includes a casing 21containing (e.g. enclosing) at least one gear train, and the compressormodule 24 includes a casing 25 located outside of the gearbox modulecasing 21. The compressor module casing 25 contains (e.g. encloses) thecompressor rotor(s) 14′ (e.g. impeller), diffuser, shroud, inlet scroll,and variable inlet guide vanes 88 (see FIG. 1) through which the aircirculates before reaching the compressor rotor(s). The compressormodule casing 25 may include a plurality of casing pieces cooperating todefine an enclosure containing the compressor 14, and/or may be definedin whole or in part by outer walls of the compressor 14. Referring toFIGS. 3-4, the compressor module casing 25 is mounted on a face of thegearbox module casing 21. In a particular embodiment, the compressormodule casing 25 and the gearbox module casing 21 are detachablyinterconnected, for example by having abutting flanges of the casings25, 21 interconnected by bolts and/or clamps or through the use of anyother appropriate type of fasteners, including, but not limited to, suchengagement members or fasteners defining a type of connection known as“quick access disconnect”. Other configurations are also possible.

Referring to FIG. 4, in a particular embodiment the communicationbetween the outlet of the compressor 14 and the inlet of the engine core12 is performed through an intake manifold 15. In a particularembodiment, the compressor rotor(s) are sized to supply engine mass flowand cabin air bleed. The intake manifold 15, which may be providedseparately from the compressor module 24, includes a branch-off port 15′for pressurized cabin bleed air.

The turbine module 28 includes a turbine module casing 29 containing(e.g. enclosing) the turbine section 18, including at least one rotorconnected to a turbine shaft 19, with respective turbine vane(s),housing(s), containment feature(s) and tie-bolt(s). The turbine modulecasing 29 is spaced from the compressor module casing 25 and alsolocated outside of the gearbox module casing 21. The turbine modulecasing 29 may include a plurality of casing pieces cooperating to definean enclosure containing the turbine section 18 and/or may be defined inwhole or in part by outer walls of the turbine section 18. The turbinemodule casing 29 is mounted on the face of the gearbox module casing 21opposite that receiving the compressor module casing 25; in a particularembodiment, the turbine module casing 29 is mounted on the forward faceof the gearbox module casing 21. In a particular embodiment, the turbinemodule casing 29 and the gearbox module casing 21 are detachablyinterconnected, for example by having abutting flanges of the casings29, 21 interconnected by bolts and/or clamps or through the use of anyother appropriate type of fasteners, including, but not limited to, suchengagement members or fasteners defining a type of connection known as“quick access disconnect”. Other configurations are also possible.

A plurality of exhaust pipes 30 provide the fluid communication betweenthe outlet of the engine core 12 (exhaust port of each engine 12′) andthe inlet of the turbine section 18. The core module 22 is mounted onthe same face of the gearbox module casing 21 as the turbine module 28;in a particular embodiment, close-coupling of the turbine module 28 tothe core module 22 helps increase (and preferably maximize) exhaust gasenergy recovery by keeping the exhaust pipes 30 between the engine core12 and the turbine section 18 as short as possible and controlling theflow area throughout. The exhaust pipes 30 become very hot during use,and accordingly appropriate materials selection and cooling isimplemented to ensure their durability.

As can be seen in FIG. 1, the turbine section 18 may include one or moreturbine stages contained in the turbine module casing. In a particularembodiment, the turbine section 18 includes a first stage turbine 26receiving the exhaust from the engine core 12, and a second stageturbine 27 receiving the exhaust from the first stage turbine 26. Thefirst stage turbine 26 is configured as a velocity turbine, also knownas an impulse turbine, and recovers the kinetic energy of the coreexhaust gas while creating minimal or no back pressure to the exhaust ofthe engine core 12. The second stage turbine 27 is configured as apressure turbine, also known as a reaction turbine, and completes therecovery of available mechanical energy from the exhaust gas. Eachturbine 26, 27 may be a centrifugal or axial device with one or morerotors having a circumferential array of radial, axial or mixed flowblades. In another embodiment, the turbine section 18 may include asingle turbine, configured either as an impulse turbine or as a pressureturbine.

A pure impulse turbine works by changing the direction of the flowwithout accelerating the flow inside the rotor; the fluid is deflectedwithout a significant pressure drop across the rotor blades. The bladesof the pure impulse turbine are designed such that in a transverse planeperpendicular to the direction of flow, the area defined between theblades is the same at the leading edges of the blades and at thetrailing edges of the blade: the flow area of the turbine is constant,and the blades are usually symmetrical about the plane of the rotatingdisc. The work of the pure impulse turbine is due only to the change ofdirection in the flow through the turbine blades. Typical pure impulseturbines include steam and hydraulic turbines.

In contrast, a reaction turbine accelerates the flow inside the rotorbut needs a static pressure drop across the rotor to enable this flowacceleration. The blades of the reaction turbine are designed such thatin a transverse plane perpendicular to the direction of flow, the areadefined between the blades is larger at the leading edges of the bladesthan at the trailing edges of the blade: the flow area of the turbinereduces along the direction of flow, and the blades are usually notsymmetrical about the plane of the rotating disc. At least part of thework of the pure reaction turbine is due to the acceleration of the flowthrough the turbine blades.

Most aeronautical turbines are not “pure impulse” or “pure reaction”,but rather operate following a mix of these two opposite butcomplementary principles—i.e. there is a pressure drop across theblades, there is some reduction of flow area of the turbine blades alongthe direction of flow, and the speed of rotation of the turbine is dueto both the acceleration and the change of direction of the flow. Thedegree of reaction of a turbine can be determined using thetemperature-based reaction ratio (equation 1) or the pressure-basedreaction ratio (equation 2), which are typically close to one another invalue for a same turbine:

$\begin{matrix}{{{Reaction}(T)} = \frac{\left( {t_{S\; 3} - t_{S\; 5}} \right)}{\left( {t_{S\; 0} - t_{S\; 5}} \right)}} & (1) \\{{{Reaction}(P)} = \frac{\left( {P_{S\; 3} - P_{S\; 5}} \right)}{\left( {P_{S\; 0} - P_{S\; 5}} \right)}} & (2)\end{matrix}$where T is temperature and P is pressure, s refers to a static port, andthe numbers refers to the location the temperature or pressure ismeasured: 0 for the inlet of the turbine vane (stator), 3 for the inletof the turbine blade (rotor) and 5 for the exit of the turbine blade(rotor); and where a pure impulse turbine would have a ratio of 0 (0%)and a pure reaction turbine would have a ratio of 1 (100%).

In a particular embodiment, the first stage turbine 26 is configured totake benefit of the kinetic energy of the pulsating flow exiting theengine core 12 while stabilizing the flow and the second stage turbine27 is configured to extract energy from the remaining pressure in theflow while expanding the flow. Accordingly, the first stage turbine 26has a smaller reaction ratio than that of the second stage turbine 27.

In a particular embodiment, the second stage turbine 27 has a reactionratio higher than 0.25; in another particular embodiment, the secondstage turbine 27 has a reaction ratio higher than 0.3; in anotherparticular embodiment, the second stage turbine 27 has a reaction ratioof about 0.5; in another particular embodiment, the second stage turbine27 has a reaction ratio higher than 0.5.

In a particular embodiment, the first stage turbine 26 has a reactionratio of at most 0.2; in another particular embodiment, the first stageturbine 26 has a reaction ratio of at most 0.15; in another particularembodiment, the first stage turbine 26 has a reaction ratio of at most0.1; in another particular embodiment, the first stage turbine 26 has areaction ratio of at most 0.05.

It is understood that any appropriate reaction ratio for the secondstage turbine 27 (included, but not limited to, any of theabove-mentioned reaction ratios) can be combined with any appropriatereaction ratio for the first stage turbine 26 (included, but not limitedto, any of the above-mentioned reaction ratios), and that these valuescan correspond to pressure-based or temperature-based ratios. Othervalues are also possible. For example, in a particular embodiment, thetwo turbines 26, 27 may have a same or similar reaction ratio; inanother embodiment, the first stage turbine 26 has a higher reactionratio than that of the second stage turbine 27. Both turbines 26, 27 maybe configured as impulse turbines, or both turbines 26, 27 may beconfigured as pressure turbines.

Still referring to FIG. 1, in the embodiment shown, the compressorrotor(s) 14′, first stage turbine rotor(s) 26′ and second stage turbinerotor(s) 27′ are connected to (e.g. rigidly connected to, integrallyformed with, attached to, or any other type of connection allowing therotors to rotate together with the shaft at a same speed) the turbineshaft 19, which extends through the gearbox module 20, parallel andradially offset from (i.e. not co-axial with) the engine shaft 16.

As can be seen in FIGS. 1 and 4, the compressor rotor(s) 14′ and turbinerotor(s) 26′, 27′ are cantilevered, i.e. the turbine shaft 19 isrotationally supported on only one side of the compressor rotor(s) 14′,and on only one side of the turbine rotors 26′, 27′. The turbine shaft19 is rotationally supported by a plurality of bearings 64 (e.g. rollingelement bearings such as oil lubricated roller bearings and oillubricated ball bearings, journal bearings) all located on a same sideof the compressor rotor(s) 14′, on a same side of the first stageturbine rotor(s) 26′, and on a same side of the second stage turbinerotor(s) 27′. In the embodiment shown, the bearings 64 are locatedbetween the compressor rotor(s) 14′ and the turbine rotors 26′, 27′ andcontained within the gearbox module casing 21, without additionalbearings being provided outside of the gearbox module 20. The rotatingassembly of the compressor module 24 and of the turbine module 28 isdynamically designed to rotate in a cantilevered manner, with thecritical modes of deflection outside of the engine's operatingconditions. Accordingly, the compressor module 24 and turbine module 28do not include bearings, and are thus not part of the bearing lubricantcirculation system 66, which is contained within the gearbox modulecasing 21. This eliminates the need to provide external lubricant (e.g.oil) feed or scavenge lines on the compressor module 24 and on theturbine module 28, which may facilitate removal of the compressor module24 and of the turbine module 28 from the remainder of the compoundengine assembly 10.

Alternately, the compressor 14 and turbine section 18 can each havetheir own dedicated shaft, for example for optimum componentperformance. In this case, the compressor shaft may also be onlysupported by bearings all located on a same side of the compressorrotor(s) 14′, for example in the gearbox module casing 21, such that thecompressor rotor(s) 14′ are supported in a cantilevered manner. Thecompressor rotor(s) 14′ is in driving engagement with the turbine shaft19 and/or the engine shaft 16, for example by having the compressorshaft mechanically linked with the turbine shaft 19 and/or the engineshaft 16 through a gear train of the gearbox module 20.

Still referring to FIG. 1, the gearbox module 20 is a combining gearboxmodule 20, including both a compounding gear train 68 and one or moreaccessory gear train(s) 70 contained in the gearbox module casing 21.The turbine shaft 19 is mechanically linked to, and in drivingengagement with, the engine shaft 16 through the compounding gear train68, such that the mechanical energy recovered by the turbine section 18is compounded with that of the engine shaft 16. In a particularembodiment, the compounding gear train 68 includes offset gears. In aparticular embodiment, the elements of the compounding gear train 68 areconfigured to define a reduction ratio allowing each module to operateat its optimum operating speed. The reduction ratio may accordinglydepend on engine sizing and/or other factors. In a particularembodiment, the reduction ratio is approximately 5:1; other values arealso possible.

In a particular embodiment, having the compressor and turbine rotors14′, 26′, 27′ on a same shaft 19 allows for the compounding gear train68 to be lighter, as the compounding gear train is sized to transmitonly the portion of the turbine power remaining after driving thecompressor 14.

It is understood that other types of gear trains are also possible,particularly, although not exclusively, for other configurations of therelative position between the modules. For example, in an alternateembodiment, the turbine section 18 and/or compressor section 14 may bepositioned such that its rotating components rotate coaxially with theengine shaft 16, and a planetary gear system may provide the mechanicallink and driving engagement between the engine shaft 16 and the shaft ofthe turbine section 18 and/or compressor section 14. Otherconfigurations are also possible.

The accessory gear train(s) 70 connect (mechanically link) one or moreaccessories 72 with the engine shaft 16 and/or the turbine shaft 19. Theaccessories 72 are mounted on the same face of the gearbox module casing21 as the compressor module 24 and may include, but are not limited to,one or any combination of the following: starter, fuel pump, oil pump,coolant pump, aircraft hydraulic pump, aircraft air conditioningcompressor, generator, alternator, permanent magnet alternator. In aparticular embodiment, the accessory gear train 70 includes an offsetgear system. Other configurations are also possible, including, but notlimited to, the combination of offset and planetary gear systems.

Referring to FIGS. 3-4, the proximity of the turbine module 28 to thecore module 22, and the gearbox module 20 located between the hot side(turbine module 28 and core module 22) and the cold side (compressormodule 24 and accessories 72) enables the delimitation of a relativelysmall fire zone, which in a particular embodiment simplifies the designof the aircraft nacelle and of the fire suppression system, improvingfire safety for the remainder of the compound engine assembly. In theembodiment shown, the compound engine assembly 10 includes acircumferential firewall 63 extending circumferentially around thegearbox module casing 21 and radially outwardly therefrom. The firewall63 is located such that the hot zone or fire zone (turbine module28/core module 22) is located on one side thereof, and the accessories72 and compressor module 24 are located on the other side thereof—i.e.the axial location of the firewall 63 is between that of the turbinemodule 28 and core module 22, and that of the accessories 72 andcompressor module 24.

Additional firewalls are provided to isolate the fuel system 13 from thehot turbine module 28 and the turbine exhaust pipes 30. In theembodiment of FIG. 3, two axial firewalls 65, 67 extend from thecircumferential firewall 63; the axial firewalls 65, 67 extend axiallyalong the core module 22, and radially outwardly therefrom. These twoaxial firewalls 65, 67 are circumferentially spaced from one anothersuch that the fuel system 13 is located therebetween; one of thefirewalls 65 may be located at or about top dead center position of therotary engines 12′. In the embodiment shown, the axial firewalls 65, 67are respectively located at or about the 12 o'clock position (top deadcenter) and the 4 o'clock position. An additional circumferentialfirewall 69 is axially spaced from the first circumferential firewall 63and extends between the axial firewalls 65, 67, circumferentially aroundpart of the core module 22, and radially outwardly from the core module22. The fuel system 13 is thus enclosed in a perimeter defined by thefirewalls 63, 65, 67, 69, which separate it from the turbine module 28,accessories 72 and compressor module 24.

In a particular embodiment, the firewalls 63, 65, 67, 69 extend radiallyoutwardly to the position of the nacelle contour, such that the nacellecooperates with the perimeter defined by the firewalls 63, 65, 67, 69 toenclose the fuel system 13 separately from the accessories 72,compressor module 24 and turbine module 28, and cooperates with thefirst circumferential firewall 63 to enclose the turbine module 28 andcore module 22 separately from the accessories 72 and compressor module24. In another embodiment, additional firewalls positioned radiallyinwardly of the nacelle contour may be provided to cooperate with thefirewalls 63, 65, 67, 69 to form the enclosure containing the fuelsystem 13 and the enclosure containing the turbine module 28 and coremodule 22 independently of the nacelle, in order to provide for smallerenclosures than the enclosures that would be defined by the nacelle.

In a particular embodiment, no electrical elements or accessories areincluded in the turbine module 28, which reduces or eliminates the riskof fire in the turbine module 28 in case of fuel leak. Sensors andelectrical elements other than those associated with the core module 22are all located on the cold side of the gearbox module 20 where thetemperature is not high enough to light a fire, and are separated fromthe hot zone by the firewall 63; the fuel system 13 is further separatedfrom the remainder of the hot zone, including the turbine module 28 andexhaust pipes 30, by the firewalls 65, 67, 69, to further minimize therisk of fire.

It is understood that in FIG. 3, the firewalls 63, 65, 67, 69 have beenschematically illustrated as transparent for clarity purposes, to avoidobstructing view of the other components of the engine 10, but that suchillustration does not imply a need for the firewalls 63, 65, 67, 69 tobe made of transparent material. The firewalls 63, 65, 67, 69 are madeof any material which is sufficiently resistant to high temperature asper current certification requirements. In a particular embodiment, thefirewalls 63, 65, 67, 69 are made of a material able to resist atemperature of 2000° F. for 5 minutes. An example of suitable materialis steel, but suitable other materials may be used.

Referring to FIG. 5, the compound engine assembly 10 is a reversed flowassembly. The compound engine assembly 10 includes an inlet duct 74having an inlet 76 communicating with ambient air outside of or aroundthe assembly 10, for example ambient air outside of a nacelle receivingthe assembly. The inlet duct 74 includes an inertial particle separator78 at its downstream end. Immediately downstream of the inertialparticle separator 78, the inlet duct communicates with a first conduit80 communicating with the compressor 14 and a second conduit 82 definingan inlet bypass duct communicating with ambient air outside of or aroundthe assembly 10, for example through communication with the exhaust duct84 (see FIG. 6) of the compound engine assembly 10. The first conduit 80defines a sharp turn with respect to the inlet duct 74 (e.g. byextending approximately perpendicular thereto), extending at asufficient angle from the inlet duct 74 such that the heavier particles(e.g. ice, sand) continue to the downwardly angled second conduit 82while the air follows the sharp turn of the first conduit 80. Thesection of the inlet duct 74 defining the inertial particle separator 78and the first and second conduits 80, 82 are sized to achieve adequateair velocities to ensure separation of the particles.

Still referring to FIG. 5, during engine operation, the ambient airpenetrates the compound engine assembly 10 through the inlet 76 of theinlet duct 74 on one end of the assembly 10, and circulates through theinlet duct 74 in a first direction across a length of the assembly 10.The air reaches the compressor 14 after having passed through theinertial particle separator 78, turned into the conduit 80, andcirculated through a filter 86. Inlet guide vanes 88 modulate the flowinto the compressor 14. The air is pressure boosted by the compressor 14and routed to the engine core 12; although not shown, the air flowbetween the compressor 14 and engine core 12 may circulate in part or inwhole through an intercooler. The engine core 12 further compresses theair. Fuel is injected in the engine core 12 and combusted, and work isextracted during the expansion cycle of the engine core 12. Exhaust fromthe engine core 12 is circulated to the turbine section 18. Work isfurther extracted by the turbines (e.g. impulse turbine, then pressureturbine) to drive the compressor 14, and the remaining useful work istransmitted to the engine shaft 16 via the gearbox module 20. Theair/gases circulation from the compressor 14 to the turbine section 18is done along a direction generally opposite of that of the aircirculation within the inlet duct 74, such that the exhaust gases exitthe turbine section 18 near the same end of the assembly 10 as the inlet76 of the inlet duct 74.

In the embodiment shown, a fraction of the turbine exhaust flow is usedfor anti-icing/de-icing of the inlet 76 of the assembly 10. The turbineexhaust communicates with a first exhaust conduit 90 communicating withthe exhaust duct 84 and with a second exhaust conduit 91 communicatingwith one or more conduits 92 located in the lip of the inlet 76, whichthen also communicate with the ambient air outside of or around theassembly 10, for example directly, through communication with theexhaust duct 84, or through communication with the second conduit (inletbypass duct) 82. A valve 94 can be provided at the entry of the secondexhaust conduit 91 to regulate the flow of exhaust air being circulatedin the lip conduit(s) 92 and/or to close the flow when de-icing is notnecessary.

In addition or in the alternative, anti-icing could be achieved with hotcoolant from a heat exchanger (cooler) 96 (see FIG. 6) of the assembly10, for example by having part of a hot coolant flow exiting the enginecore 12 circulating through a coil tube 98 disposed in the lip of theinlet 76 before being circulated to the associated heat exchanger 96.

Still referring to FIG. 5, it can be seen that the turbine shaft 19 isparallel to and radially offset from (i.e., non-coaxial to) the engineshaft 16, and that both shafts 16, 19 are radially offset from (i.e.,non-coaxial to) the inlet duct 74. In the embodiment shown, the shafts16, 19 are radially offset from a longitudinal central axis 100 of atleast part of the inlet duct 74, or of the whole inlet duct 74. The airflow within the inlet duct 74 occurs along a direction corresponding toor substantially corresponding to that of the central axis 100. It isunderstood that the central axis 100 may be a straight line (straightduct) or a curved line (curved duct e.g. single curve, S-shaped). In aparticular embodiment, the central axis 100 is parallel to the shafts16, 19. Other configurations are also possible, including, but notlimited to, the central axis 100 extending at a non-zero angle withrespect to the shafts 16, 19. In embodiments where the inlet duct 74 hasa curved shape (e.g.), an imaginary line may be defined as the straightline more closely corresponding to the curved central axis of the inletduct 74; this imaginary line may be parallel to the shafts 16, 19 orextend at a non-zero angle with respect thereto.

FIG. 6 shows an example of relative angular positions of the turbineshaft 19, the assembly inlet 76 and inlet duct 74, a lubricant (e.g.oil) heat exchanger 102 for cooling of the oil or other lubricantcirculated through the compound engine assembly 10 (e.g. to lubricatethe bearings of the shafts 16, 19 and the rotor(s) of the engine core12), and the coolant (e.g. water) heat exchanger 96 for cooling thecoolant circulated through the housing of the engine core 12. In aparticular embodiment, the layout of the compound engine assembly 10 issuitable for a compact streamlined nacelle with minimum aircraft drag.

The radial offset of the turbine shaft 19 and of the inlet duct 74 withrespect to the engine shaft 16 allows for the compressor and turbinemodules 24, 28, inlet duct 74, and heat exchangers 96, 102 to beclockable around the engine shaft 16, i.e. to be disposed in a varietyof angular positions around the engine shaft 16 to suit specificaircraft nacelle designs. For example, the configuration of FIG. 6 couldbe modified by placing the compressor and turbine modules 24, 28 closerto the nacelle exhaust, e.g. more toward the bottom of the assembly 10,to reduce or minimize the length of the exhaust duct 84 and/or exhaustconduits 90, 91 connected to the exhaust duct 84. The angular positionof the assembly inlet 76 and inlet duct 74 around the engine shaft 16can also be changed to suit specific aircraft nacelle designs. Thecoolant and lubricant heat exchangers 96, 102 can for example be locatedon the sides of the core module 22, at the top of the core module 22, orbehind the core module 22 as suitable for the particular aircraftassociated with the compound engine assembly 10 and/or to provideincreased accessibility to the heat exchangers 96, 102 and othercomponents for ease of maintenance, repair and/or replacement. Theaccessories 72 may be located all at a same angular position, andclocked around the core module 22 as required with respect to availablespace to receive the compound engine assembly 10. In a particularembodiment, locating all of the accessories 72 at a same angularposition allows for all of the accessories 72 to be accessible through asingle compartment access panel.

Referring back to FIG. 4, in a particular embodiment the compound engineassembly 10 is mounted to the aircraft through a mount cage 104including struts 106 connected to two opposed side mounts 105 attachedto the casing 21 of the gearbox module 20. In the embodiment shown, twostruts 106 are connected to each side mount 105 through an isolator 103,which may include for example a suitable elastomeric material. Thestruts 106 extending from the same mount 105 are angled with respect toone another such as to extend further apart from each other as distancefrom the mount 105 increases. The mount cage 104 includes a lowertransverse bar 106′ interconnecting the two lower struts extending fromdifferent mounts, and an upper transverse bar 106″ interconnecting thetwo upper struts extending from different mounts; the struts 106 areinterconnected by the bars 106′, 106″ at their ends opposite the mounts105, which are configured to be attached to the aircraft (e.g. to abulkhead of the aircraft). An arcuate support 107 extends under theengine 10 between the mounts 105. The struts 106 are positioned such asto avoid crossing the exhaust pipes 30. In a particular embodiment, sucha configuration avoids having any hot gas leak from the core engineexhaust pipes 30 into the turbine module 28 impinging onto the mountstructure (including isolators 103, fasteners, etc.), and thus avoidscompromises in mount structural integrity which could result from suchleaks impinging onto the mount structure.

In the embodiment shown, the mount cage 104 and the mounts are locatedout of the fire zone (turbine module 28/core module 22). The mount cage104, including the struts 106 and the transverse bars 106′, 106″, aswell as the mounts are located on the “cold side” of the gearbox modulecasing 21, and separated from the turbine module 28, core module 22 andexhaust pipes 30 by the firewall 63. The mount cage 104 is thuscompletely contained within an axial space extending axially from afirst location at the cold end of the assembly to a second location onthe gearbox module casing 21, with the turbine module 28, core module 22and exhaust pipes 30 being located outside of this axial space.Accordingly, the struts 106 are not challenged by the hot temperature ofthe turbine module 28, exhaust pipes 30 and core module 22, which mayhelp improve the structural integrity of the mount cage 104 and of itsconnection with the engine 10.

Referring to FIGS. 7-10, a compound engine assembly 210 according to analternate embodiment is shown, where elements similar to or identical tothe corresponding elements of the compound engine assembly 10 areidentified by the same reference numerals and will not be furtherdescribed herein. As shown in FIGS. 7-8, the compound engine assembly210 is configured as a reversed flow single shaft engine and includesfive (5) major modules: the core module 22, the gearbox module 20, thecold section/compressor module 24, the hot section/turbine module 28,and a reduction gearbox module 220. In the compound engine assembly 210,the rotatable load driven by the engine shaft 16 of the core module 22is a propeller 208. The engine shaft 16 is engaged to the propeller 208through the reduction gearbox module 220. The core module 12 is depictedas including three (3) rotary engines 12′, but is it understood that anyother adequate number of rotary engines or of other types of internalcombustion engines may be used.

In the embodiment shown, the reduction gearbox module 220 comprises aplanetary gearbox system; other configurations are also possible,including, but not limited to, offset gearbox and double-branch offsetgear train. Although not shown, additional accessories may bemechanically linked to and drivingly engaged to the reduction gearboxmodule.

Referring to FIG. 9, in use, the ambient air penetrates the compoundengine assembly 210 through the inlet 76 of the inlet duct 74,circulates through the inlet duct 74, through the inertial particleseparator 78, changes direction to circulate across the filter 86, inletguide vanes 88, compressor 14, optional intercooler 217 (see FIG. 10),and engine core 12. Exhaust from the engine core 12 is circulated to theturbine section 18 (which may include two turbine stages as previouslydescribed), where work is further extracted to drive the compressor. Theremaining useful work is transmitted to the engine shaft 16 via thegearbox module 20. It can be seen that a fraction of the turbine exhaustflow can be circulated to the lip conduit 92 for anti-icing of the lipof the inlet 76, as described above.

The firewall 63 extends from the gearbox module casing 21 between thefire zone (turbine module 28/core module 22) and the accessories 72 andcompressor module 24, as described above.

The compound engine assembly 210 also includes a turbine shaft 19parallel to and radially offset from (i.e., non-coaxial to) the engineshaft 16, with both shafts being radially offset from (i.e., non-coaxialto) the central axis 100 extending along the length of part of or of thewhole of the inlet duct 74. The central axis 100 may be parallel to theshafts 16, 19, may be a straight line extending at a non-zero angle withrespect to the shaft 16, 19 or may be curved (e.g. single curve,S-shaped). In embodiments where the inlet duct 74 has a curved shape animaginary line may be defined as the straight line more closelycorresponding to the curved central axis of the inlet duct 74; thisimaginary line may be parallel to the shafts 16, 19 or extend at anon-zero angle with respect thereto. The radial offset of the turbineshaft 19 and of the inlet duct 74 with respect to the engine shaft 16allows for the compressor and turbine modules 24, 28, inlet duct 74 andheat exchangers 96, 102 to be clockable around the engine shaft 16, i.e.to be disposed in a variety of angular positions around the engine shaft16 to suit specific aircraft nacelle designs, as described above.

Referring to FIG. 10, the compound engine assembly 210 also includes amount cage 104 including angled struts 106 connected to opposed sidemounts 105 attached to the casing 21, and configured such that thestruts 106 do not cross the exhaust pipes 30; transverse bars 106′, 106″respectively interconnecting the two lower struts and the two upperstruts extending from different mounts; and an arcuate support 107extending under the engine 210 between the mounts 105. An additionalarcuate support 207 may be provided under the engine 210 to support thereduction gearbox module 220, and a link 209 may extend on each side ofthe engine 210 between the two arcuate supports 107, 207; alternately,the additional support 207 and links 209 may be omitted. As describedabove, in a particular embodiment the mount cage 104 and the mounts 105are separated from the turbine module 28, core module 22 and exhaustpipes 30 by the firewall 63 (FIG. 8).

Referring to FIGS. 11-12, a compound engine assembly 310 according to analternate embodiment is shown, where elements similar to or identical tothe corresponding elements of the compound engine assemblies 10, 210 areidentified by the same reference numerals and will not be furtherdescribed herein. The compound engine assembly 310 is configured as areversed flow single shaft engine and includes four (4) major modules:the core module 22, the cold section/compressor module 24, the hotsection/turbine module 28, and the gearbox module including first andsecond sub-modules or parts 320, 320′ which cooperate to together definea module similar to the gearbox module 20 previously described. Althoughnot shown, the compound engine assembly 310 could be configured as aturboprop engine with a reduction gearbox module.

In a particular embodiment, the compound engine assembly 310 is, asidefrom its gearbox module 320, 320′ configured similarly or identically tothe compound engine assembly 10 or to the compound engine assembly 210previously described; it is accordingly understood that any element andcombination of elements of the assemblies 10, 210 as previouslydescribed, can be used in the assembly 310.

The first part 320 of the gearbox module includes a casing 321containing (e.g. enclosing) a first part 368 of the compounding geartrain (shown here as a pinion gear), and the second part 320′ of thegearbox module includes a casing 321′ containing a complementary part368′ of the compounding gear train. The two gearbox module casings 321,321′ are detachably interconnected; in the embodiment shown, the casings321, 321′ include complementary flanges 323, 323′ which are boltedtogether with a setting spacer 331 therebetween. However, any othersuitable type of connection may be used, including but not limited tothose described above.

The turbine shaft 19, to which the rotors of the turbine module 28 andof the compressor module 24 are connected to (e.g. rigidly connected to,integrally formed with, attached to, or any other type of connectionallowing the rotors to rotate together with the shaft at a same speed),extends through the second part 320′ of the gearbox module. The parts368, 368′ of the compounding gear train cooperate to mechanically linkand in drivingly engage the turbine shaft 19 to the engine shaft 16. Therotors of the turbine module 28 and of the compressor module 24 arecantilevered, and the bearings 64 supporting the turbine shaft 19 arecontained within the casing 321′ of the second part 320′ of the gearboxmodule, without additional bearings being provided outside of thegearbox module. Alternately, the turbine module 28 and of the compressormodule 24 can each have their own dedicated shaft. The compressor module24 and turbine module 28 do not include bearings, and are thus not partof the bearing lubricant circulation system, which is contained withinthe second gearbox module casing 321′.

The compressor module casing 25 is located outside of the gearbox modulecasings 321, 321′, and is mounted on a face of the second gearbox modulecasing 321′ (e.g. detachably interconnected through any suitable type ofconnection, including but not limited to those described above). Theturbine module casing 29 is also located outside of the gearbox modulecasings 321, 321′, and is mounted on the face of the second gearboxmodule casing 321′ opposite that receiving the compressor module casing25 (e.g. detachably interconnected through any suitable type ofconnection, including but not limited to those described above).

The first part 320 of the gearbox module includes one or more accessorygear train(s) (not shown) contained in the first gearbox module casing321. Accessories (not shown) are engaged mounted on a face of the firstgearbox module casing 321 on a same side of the gearbox module 320, 320′as the compressor module 25.

The separate gearbox module casings 321, 321′ may allow the turbinemodule 28, compressor module 24 and second part 320′ of the gearboxmodule to be separated from the remainder of the engine 310 whileremaining interconnected to one another to define a “turbo machinerymodule” which may be replaced, or serviced independently of theremainder of the engine 310.

In a particular embodiment, the separate gearbox module casings 321,321′ allows the second casing 321′ adjacent the turbine module 28 to bemade of material more resistant to heat than that of the first casing321, which may help minimize cooling requirements and/or thermalprotection requirement, as opposed to a single gearbox module casingcompletely made of the material of the first casing 321. In a particularembodiment, the first casing 321 is made of aluminium, and the secondcasing 321′ is made of steel.

Although not shown, the engine 310 includes mounts for engagement with amounting structure, such as a mount cage 104 as previously described. Ina particular embodiment, the mounts are connected to the first gearboxmodule casing 321.

Although examples of the compound engine assembly 10, 210, 310 have beenshown as turboshaft and turboprop engine assemblies, it is understoodthat the compound engine assemblies can be designed for other uses,including, but not limited to, to be used as an auxiliary power unit.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Modifications which fall within the scope of the present invention willbe apparent to those skilled in the art, in light of a review of thisdisclosure, and such modifications are intended to fall within theappended claims.

The invention claimed is:
 1. A compound engine assembly comprising: aninlet duct having a longitudinal central axis, an air flow within theinlet duct occurring along a direction corresponding to that of thelongitudinal central axis; a conduit communicating with the inlet duct;a compressor having an inlet in fluid communication with the conduit,the compressor including at least one compressor rotor connected to aturbine shaft; an engine core including at least one internal combustionengine in driving engagement with an engine shaft, the engine corehaving an inlet in fluid communication with an outlet of the compressor;a turbine section having an inlet in fluid communication with an outletof the engine core, the turbine section including at least one turbinerotor connected to the turbine shaft, the turbine shaft configured tocompound power with the engine shaft; wherein the turbine shaft and theengine shaft are parallel to each other; and wherein the turbine shaft,the engine shaft and the longitudinal central axis of at least part ofthe inlet duct are all radially offset from one another, the at leastpart of the inlet duct including an inlet of the inlet duct in fluidcommunication with ambient air around the compound engine; wherein thelongitudinal axis of at least part of the inlet duct is parallel to theturbine shaft and to the engine shaft, and wherein the engine assemblyis a reversed flow assembly such that the inlet of the inlet duct iscloser to the turbine section than to the compressor.
 2. The compoundengine assembly as defined in claim 1, wherein the longitudinal centralaxis of a majority of the inlet duct is radially offset from the turbineshaft and from the engine shaft.
 3. The compound engine assembly asdefined in claim 1, wherein the turbine shaft is connected to the engineshaft through a gearbox.
 4. The compound engine assembly as defined inclaim 3, wherein the gearbox is located between the compressor and theturbine section.
 5. The compound engine assembly as defined in claim 1,wherein the compressor and the turbine section are respectively providedin a compressor module casing and in a turbine module casing, thecompressor module and turbine module casings being spaced from oneanother and removable from the assembly independently from one another.6. The compound engine assembly as defined in claim 1, wherein each ofthe at least one internal combustion engine includes a rotor sealinglyand rotationally received within a respective internal cavity to providerotating chambers of variable volume in the respective internal cavity,the rotor having three apex portions separating the rotating chambersand mounted for eccentric revolutions within the respective internalcavity, the respective internal cavity having an epitrochoid shape withtwo lobes.
 7. The compound engine assembly as defined in claim 1,wherein the turbine section includes a first stage turbine having aninlet in fluid communication with the outlet of the engine core, and asecond stage turbine having an inlet in fluid communication with anoutlet of the first stage turbine.
 8. The compound engine assembly asdefined in claim 7, wherein the first stage turbine is configured as animpulse turbine with a pressure-based reaction ratio having a value ofat most 0.2, the second stage turbine having a higher reaction ratiothan that of the first stage turbine.
 9. A compound engine assemblycomprising: an inlet duct having an inlet in fluid communication withambient air around the compound engine assembly; a compressor having aninlet in fluid communication with the inlet duct; an engine coreincluding at least one internal combustion engine in driving engagementwith an engine shaft, the engine core having an inlet in fluidcommunication with an outlet of the compressor; a turbine sectionincluding a first stage turbine having an inlet in fluid communicationwith the outlet of the engine core and a second stage turbine having aninlet in fluid communication with an outlet of the first stage turbine,each of the first stage turbine and the second stage turbine includingat least one rotor connected to a turbine shaft, the turbine shaft andthe engine shaft being in driving engagement with one another; whereinthe turbine shaft and the engine shaft are parallel to each other; andwherein the turbine shaft, the engine shaft and at least part of theinlet duct are all radially offset from one another, the at least partof the inlet duct including an inlet of the inlet duct in fluidcommunication with ambient air around the compound engine assembly;wherein a longitudinal axis of at least part of the inlet duct isparallel to the turbine shaft and to the engine shaft, and wherein theengine assembly is a reversed flow assembly such that the inlet of theinlet duct is closer to the turbine section than to the compressor. 10.The compound engine assembly as defined in claim 9, wherein alongitudinal central axis of at least part of the inlet duct is radiallyoffset from the turbine shaft and from the engine shaft.
 11. Thecompound engine assembly as defined in claim 9, wherein the turbineshaft is connected to the engine shaft through a gearbox.
 12. Thecompound engine assembly as defined in claim 11, wherein the gearbox islocated between the compressor and the turbine section.
 13. The compoundengine assembly as defined in claim 9, wherein the compressor and theturbine section are respectively provided in a compressor module casingand in a turbine module casing, the compressor module and turbine modulecasings being separate from one another and removable from the assemblyindependently from one another.
 14. The compound engine assembly asdefined in claim 9, wherein each of the at least one internal combustionengine includes a rotor sealingly and rotationally received within arespective internal cavity to provide rotating chambers of variablevolume in the respective internal cavity, the rotor having three apexportions separating the rotating chambers and mounted for eccentricrevolutions within the respective internal cavity, the respectiveinternal cavity having an epitrochoid shape with two lobes.
 15. Thecompound engine assembly as defined in claim 9, wherein the first stageturbine and the second stage turbine have different reaction ratios. 16.The compound engine assembly as defined in claim 9, wherein the firststage turbine is configured as an impulse turbine with a pressure-basedreaction ratio having a value of at most 0.2, the second stage turbinehaving a higher reaction ratio than that of the first stage turbine. 17.A method of driving a rotatable load of an aircraft, the methodcomprising: directing ambient air from outside of the compound engineassembly into the compound engine assembly through an inlet duct;directing the air from the inlet duct to an inlet of a compressor;directing compressed air from an outlet of a compressor to an inlet ofat least one internal combustion engine of a compound engine assembly;driving rotation of an engine shaft with the at least one combustionengine; driving rotation of a turbine shaft of a turbine section of thecompound engine assembly by circulating an exhaust of the at least oneinternal combustion engine to an inlet of the turbine section; andcompounding power from the turbine shaft with that of the engine shaftto drive the rotatable load; wherein the turbine shaft and the engineshaft are parallel to each other and radially offset with respect toeach other; and wherein the air is circulated through the inlet ductfrom an inlet thereof along a path radially offset from the turbineshaft and from the engine shaft; wherein the path along which the air iscirculated is parallel to the turbine shaft and to the engine shaft, andwherein the engine assembly is a reversed flow assembly such that theinlet of the inlet duct is closer to the turbine section than to thecompressor.